Production, forming, bonding, joining and repair systems for composite and metal components

ABSTRACT

Method and apparatus for the production, forming bonding, joining and repair systems composite and metal components utilising at least one pressure chamber having a displaceable abutment face, wherein fluid is circulated at an elevated temperature and pressure through the pressure chamber.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a division of U.S. patent application Ser. No.10/204,938, entitled PRODUCTION, FORMING, BONDING, JOINING AND REPAIRSYSTEMS FOR COMPOSITE AND METAL COMPONENTS, filed on Sep. 19, 2002,assigned to the assignee of the present application, the disclosure ofwhich is expressly incorporated hereby by reference; U.S. applicationSer. No. 10/204,938 is related to and claims the benefit under 35 U.S.C.§119 and 35 U.S.C. §365 of International Application No. PCT/AU01/00224,filed Mar. 2, 2001.

BACKGROUND OF THE INVENTION

The present invention generally relates to production and repair systemsfor components made from material such as metal and composites. Althoughthe present invention is applicable for use in the aeronauticalindustry, and the invention will be described with respect to thisapplication, it is however to be appreciated that the present inventionalso has general industrial application.

Modern aircraft airframes can be constructed from a variety of differentmaterials. Commonly used material include aluminium and other metals,although composite materials are also now being used to produce variousaircraft components.

The high precision nature of airframe and aircraft component productionrequire the use of specialised production methods. Furthermore,specialised equipment is required to repair such airframes orcomponents. The present invention relates to a system for repairing,joining and producing composite and metal components. Although thepresent invention will be described in relation to its use in theaeronautical industry, the invention is also applicable for otherapplications.

The term “composite” is commonly used in industry to identify componentsproduced by impregnating a fibrous material with a thermoplastic orthermocuring resin to form laminates or layers.

Composites are widely used in the aerospace industry to provide aircraftcomponents such as fuselages, wings and tail fins, doors and so on. Thisis because composite components have the physical attribute of beingrelatively lightweight while at the same time having high structuralstrength in comparison to metals. Such composite components typicallyare of a sandwich construction having a honeycomb core covered by anouter and inner skin.

It is possible to repair holes passing at least partially and totallythrough such composite components. The general approach is to remove thedamaged part from the aircraft, and repair the hole by using an electricblanket with a vacuum bag. A “prepreg” formed of a layer of fibrousmaterial impregnated with uncured resin is laid over the area to berepaired. The electric blanket applies heat to that area to cure theprepreg. The vacuum bag holds the electric blanket in position over therepair area while at the same time applying a compaction force to theprepreg.

Repairs using this approach are not however always satisfactory. This isbecause the inconsistency of the heat provided by the electric blanketleads to unreliability in the curing. Also, the use of vacuum bagcompaction is not very effective in removing air from the prepreg sothat the repaired area is not necessarily void free.

The current approach in manufacturing composite components is by the useof autoclaves for curing. This limits the maximum size of the compositecomponents that can be produced by this approach to the maximumpractical size of an autoclave. Therefore, although it is possible touse composite components in the production of the fuselage of aircraftsuch as military jet fighters, it has until now been practically toodifficult to manufacture the fuselage of larger aircraft such ascommercial transport passenger aircraft using composite componentsbecause of the size restriction and cost of large autoclaves. Thefuselage would need to be made up of a number of separate fuselagesections that would subsequently need to be joined. The compositecomponents produced using the autoclave method are furthermore fullycured. This leads to poor secondary bonding performance if two suchfully cured components are subsequently joined. Such secondary bondingis not as strong as a bond where the join area and the parts are fullycured at the same time. Composites joined in this way tend to delaminateat the secondary bond and thereby fall apart.

The autoclave method is also used when producing components made frommetal where various metal sections or parts need to be bonded usingadhesives. Again, the size of the metal component that can be producedusing this method is limited by the size of the autoclave. Also, thereis also a risk of distortion being produced in the joined metalcomponent because of the need to heat and cool the entire component.

It would be advantageous to be able to provide an improved system forrepairing aircraft components made from composite, metal or othermaterials.

It would also be advantageous to be able to produce through bonding andforming aircraft components made from composite, metal or othermaterials without the need for autoclaving in the production process.

It is therefore a first object of the present invention to provide animproved method and apparatus for repairing a composite component.

With this in mind, according to one aspect of the present invention,there is provided a method of repairing a composite component having adamaged area including:

laying a composite lay-up over the damaged area;

locating at least one pressure chamber over the damaged area, thepressure chamber including a displaceable abutment face;

circulating fluid at an elevated temperature and pressure through thepressure chamber to thereby compress the lay-up between the abutmentface of the pressure chamber and the composite and elevate thetemperature thereof to effect curing of the lay-up.

The composite lay-up may be a standard laminate comprising layers offibrous material to which curing resin is applied. Alternatively, thecomposite lay-up may be a prepreg that has been preimpregnated withcuring resin.

The abutment face may be provided by a resiliently deformable membraneformed of a resiliently deformable sheet material. For example, rubbersuch as silicone or equivalent high temperature rubbers or plasticscould be utilised.

The advantage of having a pressure chamber using a resilientlydeformable membrane, is that it takes into account variations in thethickness of the composite lay-up while at the same time providing auniform pressure throughout the lay-up. This promotes the expulsion ofair from the lay-up and the expulsion of excess resin. The resultantrepaired area may therefore have little to no air bubbles providing animproved repair.

The use of circulating fluid such as water or oils to act as the heatingmechanism provides for more even heat transfer to the lay-up. A rapidheat transfer may occur through the abutment face of the pressurechamber. Furthermore, the circulation of the heated fluid allows for amore even temperature to all areas of the composite lay-up, with boththe thinner and thicker areas of the lay-up being heated at the samegeneral temperature and heat energy.

Alternatively, the abutment face could be provided by a rigid mould facesupported in floating relation to the rest of the pressure chamber. Forexample, the mould face may be supported by a peripheral resilientlydeformable or flexible flange interconnecting the mould face to the restof the pressure chamber.

The method preferably further includes cyclically varying the pressurewithin the pressure chamber. This results in vibratory pressure wavesbeing applied to the abutment face which places the composite lay-up ina vibratory environment and acts to further facilitate the removal ofair from the lay-up. The result is a repaired area having a relativelyuniform resin to fibre ratio with little to no air bubbles in the curedcomposite lay-up.

In the case where the abutment face is a resiliently deformablemembrane, a “caul” plate or to be even more accurate on the surfacecontour a separate mould may be positioned between the abutment face ofthe pressure chamber and the lay-up. During the repair procedure, thecomposite lay-up may be sandwiched between the caul plate and thecomposite component. This caul plate or mould may be shaped to provide asurface conforming with the shape of the surface of the compositecomponent once repaired. Such a caul plate can be used when it isnecessary to ensure aerodynamic smoothness of that component. Wherevibrating pressure waves are utilised, the pressure waves may help tocompact the lay-up and consolidate the lay-up into the shape of the caulplate when in position over the lay-up.

The use of an abutment face having a floating rigid mould face alsoworks in the same manner as the caul plate, with the mould faceproviding the surface to shape the composite component.

The above-noted method may be used both when the composite component hasbeen dented and when the composite component has been puncturedcompletely therethrough. In the later case however, an initial patch mayneed to be provided on the side of the puncture opposite to the side tobe repaired. This prevents the loss of resin and so on from the damagedarea as the composite component is being repaired. Therefore, once thepatch is applied to the opposite side of the puncture, the methodaccording to the present invention can be used to repair the damagedarea of the composite component. Alternatively, the patch may be appliedat the same time as the composite lay-up.

A number of different repair sequences can therefore be used:

place an initial patch on one side of the damaged area, and repair theopposing side of the damaged area using the method and apparatusaccording to the present invention;repair each side of the damaged area using the method and apparatusaccording to the present invention; orsimultaneously repair both sides of the damaged area using the methodand apparatus according to the present invention.

The most preferred repair method is to place the mould or caul plate onthe outside skin of the composite part and compact the part from itsrear face so that any imperfections in thickness are reflected on therear face.

It may also be preferable to provide a second pressure source forbalancing the pressure applied to the damaged area to be repaired. Asecond pressure vessel may be placed on the opposite side of the damagedarea of the composite component and opposite the first pressure chamber,with the damaged area being located between the first and secondpressure chambers. The second pressure chamber may also have adisplaceable abutment face. The abutment face of the second pressurechamber may be a resiliently deformable membrane. Alternatively, theabutment face may have a floating relatively rigid face shaped toconform with the shape of the composite part being repaired.

During the repair of the composite component, the second pressurechamber may apply an equal and opposite force on the composite componentrelative to the force applied by the first pressure chamber, as thefirst and second chambers may be on opposing sides of the damaged areaof the composite.

Although the pressure chamber structure needs to be of high strengthconstruction and leak proof the use of a balancing pressure allows theopposing abutment faces of the pressure chambers to cope with highforces and pressures while at the same time being of relatively lightconstruction. The abutment faces can therefore be heated and cooledquickly with reduced use of and loss of heat energy. The total heatenergy to carry out the curing, and the curing time can therefore beminimised.

The second pressure chamber can of course also act to cure a compositelay-up providing the patch for the damaged area on the opposing surfaceof the composite component.

It should however be noted that the provision of the balancing pressureis not critical to the present invention, and the method may stilloperate without this balancing force as the pressure and thus forceapplied to the composite component by the pressure chamber can berelatively low.

It should also be appreciated that a similar method could also be usedto repair a metal component using a metal patch in place of a compositelay-up.

Therefore, according to another aspect of the present invention, thereis provided a method of repairing a metal component having a damagedarea including laying a metal patch with adhesive material over thedamaged area;

locating at least one pressure chamber over the damaged area, thepressure chamber including a displaceable abutment face;circulating fluid at an elevated temperature and pressure through thepressure chamber to thereby compress and form the metal patch betweenthe abutment face of the pressure chamber and the metal component andelevate the temperature thereof to effect compression, forming, andbonding of the metal patch. The provision of a vacuum about thecomponent being repaired during the process may also be useful.

The advantages of using this repair method for metal components are thesame as when the repair method is used on composite components. Inparticular, the use of circulating fluid at an elevated temperature andpressure through the pressure chamber helps to ensure a relativelyuniform pressure on the metal patch and a relatively uniformdistribution of the temperature across the patch. This may befacilitated by the abutment face of the pressure chamber being providedby a resiliently deformable membrane.

Furthermore, the pressure may be cyclically variant within the pressurechamber. This helps to remove air bubbles from the adhesive materialbetween the metal patch and the metal component such that air voids inthe resultant bond are minimised.

Furthermore, as with the repair of composite components, it may bepossible to use a caul plate, the metal patch being sandwiched betweenthe caul plate and the metal component. It is however preferred that thesurface of the metal component over which the metal patch is to beapplied is dimpled or indented to the depth of the metal patch if a caulplate is used in the repair. This helps to ensure a smoother finish tothe repaired area.

According to a further aspect of the present invention, there isprovided an apparatus for repairing a damaged area on a componentincluding:

at least one pressure chamber having a displaceable abutment face;

fluid circulation means for circulating fluid at an elevated temperatureand pressure through said at least one pressure chamber;

location means for locating the at least one pressure chamber over thedamaged area with the abutment face of the pressure chamber locatedadjacent the damaged area.

The pressure chamber may include a housing, with the displaceableabutment face being provided on a side of the housing. The displaceableabutment face may be a resiliently deformable membrane, or may beprovided by a floating rigid mould face as previously described. Thehousing may be shaped to generally conform with the shape of thecomposite or metal component to be repaired. For example, the housingmay be generally arcuate in shape when seen in cross-section tocorrespond with the general shape of a curved face of a fuselage panel.

Alternatively, the housing may be “U” shaped in cross-section to allowfor the repair of the edge of a wing or tail fin. Other shapes are alsoenvisaged. It should be noted that the housing does not need toprecisely follow the shape of the composite or metal component, becausethe abutment face should have enough “give” to fit the shape of thecomponent.

Alternatively, the pressure chamber housing may be hinged in one or morepoints to allow the housing to be adjusted to conform more closely tothe composite or metal component.

In the above arrangements, the resiliently deformable membrane may bedirectly secured to the housing to form the pressure chamber. It howeveris alternatively envisaged that the resiliently deformable face beprovided by an internal flexible inflating bag supported by the housing.

The fluid circulation means may include a fluid pump for circulatingfluid through fluid supply lines to and from the pressure chamber, andat least one fluid reservoir for containing fluid to be circulated.

The fluid may be pressurised by compressed gas supplied to the fluidreservoir. A compressed gas source may communicate through a gas linewith the fluid reservoir. A series of fluid reservoirs may be provided,each containing fluid at different temperatures. This allows for a rapidchange in the temperature of the fluid circulated through the pressurechamber as fluid can be sourced from different fluid reservoirsdepending on the required temperature for the fluid. Greater control ofthe curing process can therefore be achieved to rapidly heat thecomposite to cure temperature then to rapidly cool the repaired and thesurrounding laminate without leaving hot spots and thus no distortionAlternatively one fluid source with a heater to heat the fluid and aheat exchanger to cool the fluid as the cycle is completed could beemployed.

The apparatus may further include vibration or impacting means tocyclically vary the pressure within the or each pressure chamber.According to the one possible arrangement, the pressure chamber mayinclude a sonic or ultrasonic vibrator secured thereto to vibrate thepressure chamber and thereby place the composite lay-up in a vibratingenvironment. Alternatively, vibration may be achieved by an interruptervalve used to periodically interrupt the flow of fluid through thepressure chamber thereby creating a pressure wave effect commonly knownas a “water hammer”. An alternative way of achieving the pressure effectis by means of a cyclic impacting device. This can for example include a“jackhammer” or “rivet gun” type device applying a cyclic impact forceon the rigid mould itself or a plate suspended within the fluid of thepressure chamber or mounted on the structure of the pressure chamber.Other means for producing cyclic variations or pressure wave effects inthe pressure within the pressure chamber are also envisaged.

The apparatus may include two said pressure chambers which may beprovided to allow for balanced pressures to be applied to the damagedarea. The location means may therefore respectively locate the pressurechambers on opposing sides of the damaged area.

The apparatus can be used in situ on an aircraft without the need toseparate the component to be repaired from the aircraft. Location meanscan therefore be selected for this purpose. For example, vacuum pads maybe used to locate the pressure chamber over the damaged area. Inaddition vacuum can be employed on the patch itself to hold theapparatus in position, compact the patch and further assist in airremoval from the repair.

Alternatively, lines may tie the pressure chamber onto the section ofthe aircraft to be repaired. It is also envisaged that a separate freestanding structure be provided to support the apparatus.

The advantage of the in-situ repair is that the composite or metalcomponents can be repaired on the aircraft on the tarmac without theremoval and consequent loss of time required to repair it off theaircraft.

The repair method according to the present invention can also be readilyadapted for the production of composite component parts.

Therefore, a second object of the present invention is to provide animproved method and apparatus for joining composite components.

With this in mind, according to yet another aspect of the presentinvention, there is provided a method of joining composite componentsincluding:

locating two separate composite components in adjacent relation, tothereby provide a join area therebetween;

applying a joining material including at least curing resin to the joinarea:

locating at least one pressure chamber over the join area, the pressurechamber including a displaceable abutment face;

circulating fluid at an elevated temperature and pressure through thepressure chamber to thereby compress the join area with the displaceableabutment face of the pressure chamber and elevate the temperaturethereof to effect curing of the join area.

The composite components may include edges located in an adjacentrelation to provide the join area therebetween. The two compositecomponents once joined may then form a single continuous panel.

Alternatively, the edge of one said composite component may be locatedadjacent a face of another said component, for example a right anglejoint. The corner areas defined by the two composite components may thenbe the join area. The composite component extending from the face of theother component could for example be a support rib on an internal panelof an airframe or component part.

The joining material may be a laminate or prepreg that is uncured andcan be melted or cured into position across the join area. Other joiningmaterials could however be used. For example, thermoplastic sheets maybe forced by the pressure and melted by the heat into position in thejoin area. Alternatively, a resin transfer mould approach may be used,with resin being drawn into the join area under vacuum to wet and jointhe components. It is also envisaged that a resin injection mouldingapproach be used where dry fibre is used to hold the load and the resinis pumped in under pressure to wet the join area.

The displaceable abutment face of the pressure chamber may be providedby a rigid mould face supported in floating relation to the rest of thepressure chamber. For example, a resiliently deformable flange mayinterconnect the rigid mould face to the rest of the pressure chamber.The rigid mould face may provide an accurate profile to shape thesurface of the composite lay-up.

The mould face has a surface which acts to form the surface of the joinarea. For example, where the composite components are being joined toform a continuous panel, the mould face may ensure that the join areahas the same aerodynamically smooth surface as the rest of the surfaceof each composite component part.

It is however also possible for the pressure chamber to have aresiliently deformable membrane for applying pressure to the join area.A caul plate or separately positioned and accurately aligned mould canbe used between the membrane and the join area if an accurate shape andsmoothness is required for the join area.

The problem of secondary bonds can be overcome if the edges of thecomposite components for producing a continuous panel are left uncuredor only partially cured. This is not readily possible when autoclavesare used to manufacture the composite components because the curing rateand temperature of cure cannot be readily controlled over the entirecomponent.

If an autoclave is used to effect a “co-cured” bond, then the parts mustbe located in the correct position by using a jig and/or other fixtures.This jig together with the parts must be transportable to allow it to bemoved into the autoclave. This means that there is a possibility ofsubsequent misalignment and distortion of the parts being held by thejig. By comparison, in the present invention, the jig and/or fixturescan be permanently fixed to a solid surface such as a concrete floor asit is not necessary for the jig to be moved. The parts can therefore belocated in precise alignment without the need for any subsequentmovement of the jig. This leads to finer manufacturing tolerances beingachievable.

In the Applicant's Australian Patent No. 697678, details of which areincorporated herein by reference, there is described a method ofmanufacturing composite components which can allow for variation in thecuring rate of the component. This can be achieved in the describedmethod by controlling the temperature and circulation flow of the fluidbeing used to cure the component. It is then therefore possible to fullycure the component except for the edges which are allowed to remainuncured or only partially cured. When composite components manufacturedin this way are joined accordingly to the present invention, theresultant join area is not a secondary bond. This is because the edgesof each component being joined are fully cured at the same time as thejoining material resulting in a full mechanical and chemical bond.

A balancing pressure may also be provided on the opposing side of thejoin area by a second pressure chamber. The second pressure chamber mayinclude an abutment face in the form of a resiliently deformablemembrane for engaging the join area.

The provision of pressure chambers on either side of the join areaapplying opposing and balanced forces thereon will facilitate minimaldistortion of the join area.

When the or each pressure chamber is located on the join area, fluid atan elevated pressure and temperature may be circulated therethrough inthe same manner as described above for the repair method and apparatus.Similarly, the pressure may also be cyclically varied to provide avibratory pressure wave to the abutment face and subsequently to thejoin area.

It is also possible to achieve differential heating or cooling of thepart being manufactured. This may be achieved by means of supplementaryfluid chambers provided, for example by separate fluid bags, bladders ortubes located adjacent to or integral with the abutment face. Fluid maybe circulated through the supplementary fluid chamber(s) at a differenttemperature to the pressure chamber. Areas of the part adjacent thesupplementary fluid chamber can therefore be cured at a different raterelative to the rest of the part. This maybe used in allowing the edgesof the part to remain uncured or partially cured. It also allows forspecific areas within the body of the part to be partially cured toallow for subsequent bonding of the components to that area.

The above described method may also be applicable for the bondingtogether of separate metal components.

Therefore, according to a further aspect of the present invention, thereis provided a method of bonding metal components including:

locating two separate metal components in adjacent relation, to therebyprovide a bond area therebetween;

applying a bonding material including at least one metal section andadhesive to the bond area;

locating at least one pressure chamber over the bond area, the pressurechamber including a displaceable abutment face;

circulating fluid at an elevated temperature and pressure through thepressure chamber to thereby compress the bond area with the displaceableabutment face of the pressure chamber and elevate the temperaturethereof to effect bonding at the bond area.

Where the two metal components are being bonded into a single continuingpanel, then the metal section may be an elongate metal strip section.However, where the edge of one metal component being bonded to and at anangle to the face of another component, then two metal angle sectionsmay be provided, one located at each corner defined between the twocomponents.

Accordingly to yet another aspect of the present invention, there isprovided an apparatus for joining components including:

at least one pressure chamber having a displaceable abutment face;

location means for locating the pressure chamber over a join areaprovided by locating two separate components in adjacent relation; andfluid circulation means for circulating fluid at an elevated pressureand temperature through the at least one pressure chamber forcompressing and effecting curing or bonding of the join area.

The apparatus may further include a said pressure chamber, the locationmeans locating the further pressure chamber on the opposing side of thejoin area for providing a balancing pressure for the other said pressurechamber.

The abutment face of one of the pressure chambers may be in the form ofa floating rigid mould face, whereas the abutment face of the otherpressure chamber may be resiliently deformable.

The fluid circulation means may be similar to that of the repairapparatus previously described. Furthermore, vibration or impactingmeans may be included for cyclically varying the pressure or producing apressure wave effect within the pressure chamber as in the repairapparatus.

The pressure chamber may include an elongate straight housing with theabutment face provided on one side thereof. Such a pressure chamber isused for straight join areas between adjacent components.

The pressure chamber could alternatively have an annular housing withthe abutment face provided on the inner or outer periphery thereof. Sucha pressure chamber can be used for joining fuselage sections foraircraft or hull sections of boats, submarines, or other largestructures.

The joining method and apparatus according to the present invention canbe used to join together sections of an entire fuselage of an aircraftformed of a number of separate composite or metal panels. The proposedinvention is of course also applicable for producing other largestructures from composite or metal components that cannot be readilyconstructed in one piece.

The present invention may also be adapted for bonding metal sectionssuch as “top hat” and “T” sections to metal panels and the sections thata person skilled in the art would be familiar with. Top hat sections areused to support and reinforce metal panels of an airframe and arenormally secured to the metal panel using rivets. Elaborate jigs arerequired to hold the panel and top hat sections in position prior tosecuring. Alternatively, the top hat sections are bonded to the metalpanel using adhesive within an autoclave. It has however been found thatthe top hat sections tend to “flatten” somewhat due to the highpressures applied to them within the autoclave.

It is therefore another object of the present invention to provide animproved method and apparatus for bonding a top hat and other, sectionsto a metal panel.

With this in mind according to a further aspect of the presentinvention, there is provided a method of bonding a section to a metalpanel, the section having a centre portion and opposing side flangeportions, the method including locating the section adjacent the metalpanel with the flange portion thereof immediately adjacent the metalpanel, with adhesive material being applied between the metal panel andthe flange portions, locating a respective pressure chamber over eachsaid flange portion, each pressure chamber including a displaceableabutment face;

circulating fluid at an elevated temperature and pressure through eachpressure chamber to thereby compress the flange portions with thedisplaceable abutment faces of the respective pressure chamber andelevate the temperature thereof to effect bonding of the flange portionsto the metal panel.

The temperature and pressure within each pressure chamber may be thesame. This minimises the possibility of distortion to the section.Pressure vibration may also be provided by the pressure chambers asdescribed previously.

The opposing side of the metal panel may rest on a support jig.Alternatively, the opposing side of the metal panel may be supported ona floating mould similar to that shown in Australian Patent 697678.

An opposing pressure balancing the pressure may be applied to the flangeportions by the jig or floating mould. More particularly, the balancingpressure may be provided by another pair of pressure chambers similar tothat applying pressure to the flange portions. It is also possible forthe opposing pressure chambers to apply a pressure directly to theopposing face of the metal without the need of an intermediate jig orfloating mould.

The section may be a top hat section having a centre channel portion orother section that a person skilled in the art would be familiar with.Alternatively, the section may be a “T” section having a central webportion.

According to yet another aspect of the present invention, there isprovided an apparatus for bonding a section to a metal panel, thesection having a centre portion and opposing side flange portions, theapparatus including;

a pair of elongate pressure chambers located in a spaced apart parallelrelation, each pressure chamber having a displaceable abutment face;

the central section of the section being locatable between the pressurechambers, with the abutment face of each pressure chamber beingrespectively locatable over a said flange portion of the section, andfluid circulation means for circulating fluid at an elevated pressureand temperature through each pressure chamber for compressing andeffecting bonding of the flange portions to the metal panel.

The apparatus may be formed of a rigid elongate housing having a centralchannel sized to accommodate therein the central portion of the section,and opposing side channels providing the walls of the two pressurechambers. The abutment face of each pressure chamber may be provided bya resiliently deformable membrane.

The fluid circulation means may communicate at the same time with bothpressure chambers so that the fluid temperature and pressure in eachpressure chamber may be the same.

The housing may be flat or may be curved depending on the intended shapeof the metal panel and cooperating section.

A pair of opposing apparatus according to the present invention may beprovided, the metal panel and sections being locatable between the twoapparatus during the bonding process. The two apparatus may have acooperating curve so that the panel may be held therebetween in a curvedmanner to correspond to the final profile of the metal panel. Theapparatus can therefore act as a jig for the panel.

The present invention is also applicable for the production of metalcomposites such as honeycomb metal panels or metal panels formed frombonded layers of metal.

Therefore, according to a further aspect of the present invention, thereis provided a method of producing a metal composite component including;

locating metal composite material together with adhesive materialtherebetween between opposing pressure chambers, each pressure chamberhaving a displaceable abutment face;

circulating fluid at an elevated temperature and pressure through eachpressure chamber to thereby bend, form, and compress the metal compositematerial between the opposing abutment faces and elevate the temperatureof the metal composite material to effect bonding of the adhesivematerial to the metal.

An air breather cloth may be located between each abutment face and thecomposite material to facilitate the release of air therebetween.

One abutment face may be in the form of a floating pressure plate whichmay be flat or may be provided with a curved on profiled surface.Alternatively, the abutment face may be provided by a floating mould.The opposing abutment face may be provided by a resiliently deformablemembrane.

The method may further include applying cyclic pressure variationswithin each pressure chamber to help form/bend the metal composite tothe desired profile of configuration.

The metal composite may for example include a central metal honeycombcore, with metal skins being provided on opposing faces thereof. Anadhesive material may be provided between each metal skin and thecentral honeycomb cone. Each metal skin may include a plurality ofbleeder holes for allowing excess adhesive and air to be released fromthe composite.

Alternatively, the metal composite may include a plurality of metalsheet layers between which is provided sheets of adhesive material, themethod bonding and forming the metal sheets into the desired profile.The metal sheets may include perforations to allow the release of airand adhesive material.

Vacuum may assist in the consolidation of the stacked material howeverthis is not necessary for the process to operate.

It will be convenient to further describe the invention with referenceto the accompanying drawings which illustrate preferred embodiments ofthe present invention. Other embodiments are possible, and consequentlythe particularity of the accompanying drawings is not to be understoodas superceding the generality of the preceding description of theinvention.

IN THE DRAWINGS

FIG. 1 is a schematic view showing a first preferred embodiment of themethod according to the present invention used to repair a compositecomponent;

FIG. 2 is a schematic view showing a second preferred embodiment of themethod according to the present invention used to repair a compositecomponent;

FIG. 3 is a schematic view showing a third preferred embodiment of themethod according to the present invention used to repair a compositecomponent;

FIG. 4 is a schematic view showing the fluid circulation circuit for theapparatus according to the present invention;

FIG. 5 is a schematic view of a preferred arrangement for joiningfuselage sections of an aircraft according to a fourth exampleembodiment of the present invention;

FIGS. 6 a and b are schematic views showing the construction sequence ofan aircraft wing according to the present invention;

FIG. 7 is a cross-sectional view of a lateral joint of composite panelsusing the method according to a fifth preferred embodiment of thepresent invention;

FIG. 8 is a cross-sectional view of the lateral joint of FIG. 7 whencompleted;

FIG. 9 is a cross-sectional view of a lateral joint of metal panelsaccording to a sixth preferred embodiment of the present invention;

FIG. 10 is a cross-sectional view of the lateral joint of FIG. 9 whencompleted;

FIG. 11 is a cross-sectional view of a completed lateral joint for acloseout using the method and apparatus shown in FIG. 9;

FIG. 12 is a schematic cross-sectional view showing the bonding sequenceof a top hat section to a metal panel according to a seventh preferredembodiment of the present invention;

FIG. 13 is a schematic cross-sectional view showing the bonding of a tophat section to a metal panel according to an eighth preferred embodimentof the present invention;

FIG. 14 is a schematic cross-sectional view showing the bonding of a tophat section to a metal panel according to a ninth preferred embodimentof the present invention;

FIG. 15 is a schematic side view of the apparatus shown in FIG. 14 usedto bond a top hat section to a curved panel;

FIG. 16 is a schematic cross-sectional view of a tenth preferredembodiment of the present invention used to produce a metal compositepanel having a metal honeycomb core;

FIG. 17 is a cross-sectional view of the panel produced by the methodshown in FIG. 16;

FIG. 18 is a schematic cross-sectional view of an eleventh preferredembodiment of the present invention used to produce a metal compositepanel formed of layers of metal sheeting; and

FIG. 19 is a cross-sectional view of the panel produced by the methodshown in FIG. 18;

FIG. 20 is a plan view of a perforated metal panel for the metalcomposite panel of FIGS. 18 and 19; and

FIG. 21 is a plan view of an adhesive layer for the metal compositepanel of FIGS. 18 and 19.

The composite component repair system, the composite component joiningsystem, the metal component joining system and the metal composite panelproduction system all utilise the same common features. The followingdescription therefore relates to all of these systems unless otherwisespecified. It should also be noted that corresponding features indifferent embodiments of the invention are generally provided with thesame reference numeral for clarity reasons.

Referring initially to FIG. 1, there is shown a partial cross-section ofa composite component 1 having a damaged area 2 which is to be repaired.A first pressure chamber 3 is provided on one side of the damaged area2, with a second pressure chamber 4 being located on the opposite sideof the composite part 1 adjacent the damaged area 2.

It should be appreciated that where the present invention is used tojoin two composite or metal components 1, then the join area would belocated between the first and second pressure chambers 3, 4. Therefore,the description of this and the following example embodiments of theinvention are also applicable for the joining method. The first pressurechamber 3 includes a displaceable abutment face 5 in the form of aresiliently deformable membrane 5. This membrane may be made of rubbersuch as silicone or other high temperature plastics or materials. Thesecond pressure chamber 4 similarly includes a displaceable abutmentface 6 in the form of a resiliently formable membrane. A compositelay-up 7 formed of layers of laminate material impregnated with curingresin is located on one side of the composite component 1 to bridge thedamaged area 2 (or the join area). This composite layer 7 acts as a“patch” for the repair. The damage in the form of a gap 9 provided inthe skin 10 of the composite part is on the opposite side to thecomposite lay-up 7 providing the patch. A repair composite lay-up 8 isplaced over the gap 9 prior to the repair process. This furthercomposite lay-up 8 is also impregnated with curing resin. It should benoted that the composite component 1 is formed of a sandwich of twoouter skins 10, 11 between which is provided a core of honeycombmaterial 12. The repair composite lay-up 8 includes sufficient materialto both cover the gap 8 and fill any cavity provided within the core 12of the composite component 1.

During operation of the method according to the present invention, theabutment faces 5, 6 of each pressure vessel 3, 4 are forced against therespective opposing sides of the damaged area 2. Fluid at an elevatedtemperature and pressure is then circulated through each pressurechamber 3, 4 to thereby compress the respective composite lay-ups 7, 8and effect curing of the resin within each lay-up. A peel ply andbreather cloth (not shown) can also be placed between the lay-ups 7, 8to assist in the separation of the abutment face 6, 5 of each pressurevessel from the lay-up 7, 8 as well as to allow for the removal ofexcess curing resin therefrom.

FIG. 2 shows a similar arrangement to FIG. 1 except that a caul plate oraccurate mould 15 is provided between the abutment face 5 of the firstpressure chamber 3 and the repair composite lay-up 8. The caul plate 15can be made of a relatively soft metal which can be preshaped into thedesired form of the composite component 1 when repaired. The caul plate15 will ensure that the surface 10 of the repaired composite component 1has the desired shape when repaired. This is important where componentsurface 10 is required to be smooth when repaired, for example, inaircraft or boat applications and the pressure chamber can then workmost effectively in compacting the part to the desired shape against therigid mould. As the mould is or can be separate to the pressure vesselit can be positioned first accurately and rigidly to a framework to holdthe part in accurate alignment before the heat and pressure is appliedby the pressure vessels to cure the part. Thus this ensures the mostaccurate alignment possible of the mould the part and thus the curedcomponent.

FIG. 3 shows another possible embodiment of the present invention whichis also similar to the arrangement shown in FIGS. 1 and 2. The principledifference is that the first pressure chamber 3 has an abutment face 15in the form of a floating mould face 17 joined by a resilientlydeformable flange 18 to the rest of the first pressure chamber 3. Thefloating mould face 17 is relatively rigid and acts in the same manneras the caul/mould plate 15 of FIG. 2 by ensuring that the repairedcomposite component 1 has a required profile and smoothness.

In all of the above arrangements, a cyclically varying pressure wave isprovided in the fluid within both the first and second pressure chambers3, 4 during the repair (or joining) procedure to facilitate the removalof air from the composite lay-ups 7,8. Furthermore, the use of twopressure chambers provide a balanced force on the composite component 1.

FIG. 4 shows the various components of the fluid circulation system 20of the first pressure chamber 3. It should also be noted that the fluidcirculation system 20 is also connected to the second pressure chamber 4even though this is not shown in FIG. 4.

This system 20 includes three fluid reservoirs 21, 22, 23 containingfluid at different temperatures. Therefore, fluid at the highesttemperature is held in the first fluid reservoir 21, fluid at anintermediate temperature is held in the second fluid reservoir 22, whilecold fluid 23 is contained in the third fluid reservoir 23. Each of thefluid reservoirs 21, 22, 23 are pressurised by compressed air from acompressed air source 24. The compressed air is delivered through acompressed air line 25 to a conduit 26 interconnecting the upper volumesof the first and second fluid reservoirs 21, 22. A similar conduit 27interconnects the upper volume of the second and third fluid reservoirs22, 23. The same air pressure is therefore applied to the fluid in eachof the fluid reservoirs. A fluid supply line 28 supplies fluid to thepressure chamber 3, and a fluid return line 29 returns fluid from thepressure chamber 3 back to the fluid reservoirs. A fluid pump 30 isprovided on the vacuum side of the pressure chamber 3.

A vibrator 31 is attached to the pressure chamber 3. This vibrator 31induces cyclically varying pressure waves within the pressure chamber 3to therefore place the composite lay-ups 7, 8 in a vibratoryenvironment. This acts to facilitate the removal of air bubbles frombetween the layers of laminate in the composite lay-ups 7, 8. This maybe further facilitated by a balance of vibration forces from top andbottom and or with balanced frequency so that the top and bottomvibrations occur at exactly the same moment and frequency and amplitude.If this approach is not followed then only one vibration source ispractical and this should emanate from the flexible membrane sideopposite the accurate surface to be created or formed.

FIG. 5 shows the present invention adapted for joining together fuselagesections 50 for an aircraft. These fuselage sections 50 may themselvesbe formed of a series of smaller composite panels joined by the methodaccording to the present invention. Two fuselage sections 50, eachhaving a circular cross-section are supported on a jig and are boughttogether to provide a join area therebetween. The first pressure chamber3 is in the form of a segmented annular ring, with the abutment face 5of the first pressure chamber 3 being located at the inner peripherythereof. The pressure chamber may also be comprised of a separate bag orinflatable bladder that can be positioned separately to the supportingframework. The first pressure chamber 3 totally encircles the fuselagesections 50. The second pressure chamber 4 is placed within the confinesof the fuselage section 50. This second pressure chamber 4 is also inthe form of an annular ring having an abutment face 6 provided on theouter periphery thereof. This fuselage joining arrangement otherwiseoperates in an identical fashion to the previously described repair andjoining methods.

FIGS. 6 a and b shows the construction sequence of an aircraft wing 52according to the present invention. The wing 52 is constructed byinitially laying a top wing skin 54 within a mould (not shown) definingthe profile of the top wing skin 54. The top wing skin portion of thewing 52 is constructed first because of the higher loads applied to thetop of the wing 52 during flight. Wing spars 55, 56 are joined to thetop wing skin 54 using joins 60 produced by the method according to thepresent invention. This method will subsequently be described in moredetail. The bottom wing skin 53 while preferably retained in the mouldis then laid over the rest of the wing 52 and joined to the wing spars55, 56 by further joins 62 produced according to the present invention.Join material 64 for producing the join 62 is then initially secured toopposing sides of the top of each wing spar 55, 56 prior to the bottomwing skin 53 being laid on top. Bladders 66, 68 are located in thehollow within the wing 52 for providing the pressure chambers accordingto the present invention. Heated pressurised fluid is circulated throughthe bladders 66, 68 to complete the bonding of the joins 60, 62 of thewing spars 55, 56 to the wing skins 52, 53. This may be accompanied bythe pressurisation, at the same time and pressure of the fluid in thetop and bottom moulds to minimise the distortion forces on the moulds.The bladders 66, 68 can be subsequently removed from the wing 52 throughaccess holes after the joins have been formed.

The bladders 66, 68 can fill the entire volume within the wing 52between the wing spars 55, 56 as shown in FIG. 66. It is however alsoenvisaged that smaller bladders/tubes or other fluid chambers be usedsupported by a jig, within that volume. The jigs hold the fluid chamberadjacent the area to be joined. For example, elongate fluid chambers maybe located by means of a jig located within the volume of the wing sparadjacent each of the joins 60, 62. The use of such an arrangement allowsthe fluid chambers to be readily removed from the wing 52 after use. Thewing skin 52, 53 and wing spars 55, 56 can be made of compositematerial. FIG. 7 shows the method and apparatus for joining compositecomponents in a lateral joint of the type shown in FIG. 6 a and b. Thisjoining method also has general industrial application and willtherefore be described for such general applications.

FIG. 7 shows a composite panel 70 supported on a jig 72. Alternatively,it is possible to support the composite panel 70 on a floating mould ofthe type described in Australian Patent 697678. A second compositecomponent 74 can be joined to the first composite panel 70 by applying acomposite joint at each of the corners defined between the compositepanels 70, 74.

The apparatus 75 according to the present invention includes an elongatehousing 76 extending along the length of the joint, the housing 76supports a resiliently deformable membrane 78 to thereby define apressure chamber 80 therein. A similar pressure chamber 80 is providedon the opposite side of the panel 74 so that the pressures applied tothe panel 74 are balanced during the curing of the joint. This ensuresthat there is no movement of the panel 74 while the pressure is beingapplied thereto. It is however to be appreciated that one of thepressure chambers 80 could be replaced by a jig located on the opposingside of the panel 74 to the other pressure chamber 80.

The composite joint 90 is prepared by initially applying a fillet ofresin 82 in the corner to hold the lateral panel 74 in position. Thisresin fillet 82 also provides a curve upon which layers of resin wettedcloth or prepreg cloth 84 can be placed. This curve minimises thepossibility of wrinkles being formed in the final composite joint. Overthe prepreg cloth or sheets of dried fabric wetted with resin 84 isprovided a peel ply or release cloth 86. Finally, a breather cloth 88 isprovided over the release cloth. The purpose of the peel ply 86 andbreather cloth 88 have been previously discussed and will not berepeated here. Once the pressure chambers 80 are located over thevarious layers, fluid at high temperatures and under pressure is passedthrough the pressure chambers 80 to thereby heat and compress the jointmaterial. Furthermore, the pressure is cyclically varied within eachpressure chamber. By having both pressure chambers 80 supplied from thesame fluid source, this balances the pressures across the lateral panel74 and there is no force acting to displace the panel 74.

The joint material can therefore be heated and cured by the apparatusaccording to the present invention.

The final composite joint 90 is shown in FIG. 8 which shows that joint90 compacted against the corner provided between the composite panel 70,74. The various layers of prepreg may be laid such that the compositejoint 90 is thicker at the corner where the stresses are likely to behigher than at the peripheral edges of the joint 90 where there is lessstress.

The method and apparatus according to the present invention can also beused to join metal components. The apparatus 75 used to join metalcomponents is identical to the apparatus used to join compositecomponents as shown in FIG. 7. The same reference numerals are thereforeused for corresponding features. The main difference is the differentmaterials used for providing a lateral joint between two metal panels92, 94. The joint material includes a metal angle plate 96 with adhesivematerial being provided between the angle plate 96 and the metal panels92, 94. A breather cloth 88 is provided over the metal angle plate 96 toenable the release of any air bubbles trapped in the adhesive material.The pressure chambers 80 act to bond the metal angle plate 96 to thepanels 92, 94 to thereby complete the joint. The final joint is shown inFIG. 10 with the metal angle plate 96 being adhesively bonded to thepanels 92, 94. This method of bonding the angle plate 96 is more energyefficient and less time consuming than conventional methods which wouldrequire the complete assembly to be located within an autoclave to curethe adhesive. The need to heat and subsequently cool the entire assemblycan also lead to the potential of distortion in the joint or the metalcomponent.

The method of forming a joint between metal components as shown in FIG.9 can also be used to bond a closeout arrangement as shown in FIG. 11.Such arrangements are formed from honeycomb or other composite panelswhere the peripheral edge of the panel 98 is necked down to provide athin section 100 about its periphery. A metal angle plate 96 can beapplied to the join area between the peripheral edge 100 of the panels.It is to be appreciated that a composite joint as shown in FIG. 10 couldalso be used to join the panels.

Another form of construction used in the aircraft industry is metalpanels reinforced with top hat and “T” or other sections. These sectionsare at present generally riveted to the metal panel. The manufacturingprocess requires elaborate jigs to hold the panels and sections inposition while the holes are drilled and the rivets installed. This canbe a very time consuming process. The present invention can howeverallow such top hat sections to be rapidly secured to a metal panel. FIG.12 shows from left to right the various steps involved in securing a tophat section 102 to a metal panel 104. The metal panel 104 can besupported on a jig 106 or a floating mould of the type shown inAustralian Patent 697678.

The top hat section 102 includes a central channel portion 106 andopposing peripheral flange portions 108. The apparatus 109 according tothe present invention includes an elongate housing 110 having a centrechannel portion 112 of a size sufficient to accommodate the centrechannel portion 107 of the top hat section 106, and opposing sidechannels 115, each side channel supporting a resiliently deformablemembrane 117. A pressure chamber 118 is therefore provided within eachof the side channels 115.

An adhesive 119 is placed between the metal panel 114 and the flangeportions 108 of the top hat section 102 as shown on the extreme left ofFIG. 12. The top hat section 102 is placed against the metal panel 104with the two pressure chambers 118 being respectively applied over theflange portions 108 thereof. The circulation of heated fluid underpressure cures the adhesive 119 such that the top hat section 102 can bebonded to the metal panel 104. Because no pressure is applied to thecentral channel portion 107 of the top hat section 102, there is nodistortion of the top hat section 102 during the curing process.

The final secured top hat section 102 is shown on the extreme right ofFIG. 12.

FIG. 13 shows a variation of the method shown in FIG. 12 wherecorresponding apparatus 109 are provided on opposing sides of the jig106 to balance the pressure applied by the pressure chambers 118 on theflange portions of the top hat section 102.

FIG. 14 is a further variation of the method of FIG. 13 except that thejig 106 has been removed. This is because the opposing pressure chambers118 provide a balance force on the metal panel 106 eliminating the needfor a jig to support the metal panel. Furthermore, the housing 110 ofthe apparatus 109 can be curved to follow the general curvature andprofile of the top hat section and thus form the metal panel 106 to thetop hat section, when, for example, forming part of the fuselage of anaircraft. Opposing curved housings 110 can be provided on opposing sidesof a metal panel so that, as well as acting like a jig by holding themetal panel in position, they enable the top hat section 102 to besecured thereto and thus form the metal panel to the top hat section.This arrangement, which is schematically shown in FIG. 15, eliminatesthe need for any elaborate jigs to hold the metal panel 104 in itsrequired curved shape while the top hat sections 102 are being securedthereto.

It is also possible to use the present invention to produce metalcomposite panels. Referring to FIG. 16, the various components of thepanel to be produced is located between two pressure chambers (notshown) according to the present invention. One of the pressure chambersincludes a floating pressure plate 120, the other pressure chambersincluding a flexible bladder 122. The components of the panel to beconstructed includes a central honeycomb panel core 124 located betweentwo metal skins 126. The completed panel 121 is shown in FIG. 17.Adhesive material is sprayed or rolled on to the metal skin 126 and themethod according to the present invention acts to bond the metal skin tothe honeycomb core 124. A breather cloth 127 is also provided betweenthe flexible bladder and one of the metal skins 126 to allow air toescape therefrom. Furthermore, the method according to the presentinvention can act to form the metal panel 121 into the required shape.For example, the panel 121 shown in FIG. 17 has been formed in to acomplex curve. This is assisted by cyclically varying the pressurewithin the pressure chambers during the forming process. The metal skins126 are provided with perforations to allow air to be released fromwithin the core of the panel 121 during the forming process. The methodaccording to the present invention can also neck together and bond theperipheral edges of the metal skin 126.

A similar procedure can be used to produce a composite panel 130 formedfrom layers of perforated metal sheeting 132 between which is providedadhesive layers 134. The metal panels may be perforated with multipleholes 133 as shown in FIG. 20. The adhesive layer may include pathways136 for releasing air therefrom and be of firm or resilient constructionat room temperature to ensure that air release occurs during compactionand forming of the metal panels. Once compaction and forming arecomplete the parts are heated and the adhesive melts to bond the skinstogether to form a panel. Also, a reinforcement material such as kevlarcan also be included within the adhesive layer 134.

It is also possible to use the method and apparatus according to thepresent invention to provide for other manufacturing procedures such asthe super plastic deformation of metals such as aluminium sheets. Attemperatures of about 250 degrees Celcius, such material can be readilydeformed. This can be achieved by locating a sheet to be deformedbetween opposing pressure chambers and circulating fluid at or above theabove noted temperature while also providing a pressure wave effect or acyclic pressure variation in the circulating fluid. It should be notedthat different types of metals can be employed for various layers of themetal panel, for example, the outer skin may be thin titanium to resistcorrosion and erosion and the inner layers may be lithium aluminium forlight weight and ease of construction.

1-57. (canceled)
 58. A method of joining composite components including:locating portions of two separate composite components in adjacent oroverlapping relation, to thereby provide a join area therebetween; eachsaid composite component being substantially fully cured, with each saidportion located within the join area being uncured or partially cured;locating at least one pressure chamber over the join area, the pressurechamber including a displaceable abutment face; circulating fluid at anelevated temperature and pressure through the pressure chamber tothereby compress the join area with the displaceable abutment face ofthe pressure chamber and elevate the temperature thereof in a controlledmanner over time to effect curing of the join area.
 59. A methodaccording to claim 58, including cyclically varying the pressure withinthe pressure chamber.
 60. A method according to claim 58, wherein thedisplaceable abutment face is provided by a resiliently deformablemembrane formed of a resiliently deformable sheet material.
 61. A methodaccording to claim 58, wherein the displaceable abutment face isprovided by a separate flexible inflating bag.
 62. A method according toclaim 58, including locating a second pressure chamber having adisplaceable abutment face at an opposing side of the join area.
 63. Amethod according to claim 62, wherein the second chamber has fluidpressure and the fluid pressure in each pressure chamber is at leastsubstantially identical.
 64. A method according to claim 58, includinglocating the components to be joined in alignment using at least onefixed jig, the at least one pressure chamber being moveable for locationover the join area.
 65. A method according to claim 58, furthercomprising applying a joining material including at least prepreg orcuring resin to the join area.
 66. A method of bonding metal componentsincluding: locating two separate metal components in adjacent relation,to thereby provide a bond area therebetween; applying a bonding materialincluding at least one metal section and adhesive to the bond area;locating at least one pressure chamber over the bond area, the pressurechamber including a displaceable abutment face; circulating fluid at anelevated temperature and pressure through the pressure chamber tothereby compress the bond area with the displaceable abutment face ofthe pressure chamber and elevate the temperature thereof to effectbonding at the bond area.
 67. A method of bonding a section to a metalpanel, the section having a centre portion and opposing side flangeportions, the method including locating the section adjacent the metalpanel with the flange portion thereof immediately adjacent the metalpanel, with adhesive material being applied between the metal panel andthe flange portions, locating a respective pressure chamber over eachsaid flange portion, each pressure chamber including a displaceableabutment face; circulating fluid at an elevated temperature and pressurethrough each pressure chamber to thereby compress the flange portionswith the displaceable abutment faces of the respective pressure chamberand elevate the temperature thereof to effect bonding of the flangeportions to the metal panel.
 68. A method according to claim 67, thesection being a top hat section.
 69. A method according to claim 67, thesection being a T section.
 70. A method of producing a metal componentincluding: locating at least one metal sheet to be deformed betweenopposing pressure chambers, each pressure chamber having a displaceableabutment face; circulating fluid at a temperature at which the materialat least one metal sheet can be readily deformed; further providing apressure wave effect or cyclic pressure variation in the circulatingfluid; such that the at least one metal sheet undergoes super plasticdeformation.
 71. A method according to claim 70, wherein a plurality oflayered metal sheets are located between the pressure chambers.
 72. Amethod according to claim 70, wherein the material of the metal sheet(s)includes aluminium, lithium-aluminium alloy and titanium.
 73. A methodof joining composite components including: locating portions of twoseparate composite components in adjacent or overlapping relation, tothereby provide a join area therebetween, each said composite componentbeing substantially fully cured with each said portion located withinthe join area being uncured or partially cured; locating at least onepressure chamber over the join area, the pressure chamber including adisplaceable abutment face; circulating fluid at an elevated pressurethrough the pressure chamber to thereby compress the join area with thedisplaceable abutment face of the pressure chamber and controlling thetemperature thereof by supplying the circulating fluid at differenttemperatures to the pressure chamber to effect curing and subsequentcooling of the join area.
 74. A method according to claim 73 includingapplying a joining material to the join area.
 75. A method according toclaim 73 wherein the circulating fluid is contained in a plurality offluid reservoirs, each reservoir containing fluid at a differenttemperature, the method including supplying the circulating fluid fromdifferent reservoirs, to thereby allow for a rapid and controlled changein the temperature both up and down through the curing cycle of thefluid circulated through the pressure chamber.
 76. A method of joiningthermoplastic composite components including: locating two separatecomposite components in adjacent or overlapping relation, to therebyprovide a join area therebetween; locating at least one pressure chamberover the join area, the pressure chamber including a displaceableabutment face; circulating fluid to an elevated pressure through thepressure chamber to thereby melt compress and form the join area withthe displaceable abutment face of the pressure chamber and controllingthe temperature thereof by supplying the circulating fluid at differenttemperatures to the pressure chamber to effect melting and subsequentcooling of the join area.
 77. A method according to claim 76 furthercomprising applying a joining material including at least onethermoplastic sheet to the join area.
 78. A method according to claim 76including cyclically varying the pressure with the pressure chamber.